1. Field of the Invention
The present invention relates to gas turbine engines, and more specifically to cooling of turbine airfoils.
2. Description of the Related Art Including Information Disclosed under 37 CFR 1.97 and 1.98
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. The hot gases flow downstream through turbine stages which extract energy therefrom for powering the compressor and producing useful work, such as powering a fan for propelling an aircraft in flight.
A high pressure turbine is disposed immediately downstream from the combustor and receives the hottest combustion gases from the combustor. The first stage turbine rotor blades have hollow airfoils which are supplied with a portion of air bled from the compressor for use as a coolant in removing heat from the blades during operation.
Each airfoil includes pressure and suction sidewalls joined together at opposite leading and trailing edges, and extending from root to tip. A platform is disposed at the airfoil root and defines a portion of the radially inner flow path for the combustion gases. And, a dovetail is integrally jointed to the platform for mounting the individual blades in corresponding dovetail slots in the perimeter of a rotor disk.
Since the airfoil leading edge first engages the hot combustion gases, it requires substantial cooling for obtaining a useful blade life. Heat load from the combustion gases varies around the outer surface of the airfoil from the leading to trailing edges, and along the pressure and suction sidewalls. Various cooling circuits are provided inside the airfoil for cooling the different portions thereof. The different portions of the airfoil therefore operate at different temperatures, which introduce thermal stress therein that affect low cycle fatigue life of the blade.
Airfoil cooling may be affected using convection cooling, film cooling, or impingement cooling, or combinations thereof. The leading edge of a first stage turbine airfoil typically includes several rows or columns of film cooling holes fed by a common leading edge flow chamber or channel. Other film cooling holes and trailing edge holes may be fed by corresponding internal channels, such as multi-pass serpentine cooling channels.
The airfoil may include additional film cooling holes disposed in either sidewall (pressure side or suction side) downstream of the leading edge, which are typically referred to as gill holes. Since the gill holes are typically provided with a common source of coolant inside the airfoil, and the pressure of the combustion gases outside of the airfoil varies, backflow margin across the gill holes may vary on opposite sides of the airfoil.
Backflow margin is defined as the pressure of the coolant inside the airfoil divided by the local pressure of the combustion gases outside the airfoil as experienced by each of the gill holes. Sufficient backflow margin must be maintained to prevent ingestion of the hot combustion gases into the airfoil, and ensure continuous discharge of the coolant through the gill holes.
Since the minimum required backflow margin must be set at the airfoil leading edge pressure sidewall, the backflow margin on the lower suction sidewall of the airfoil may be undesirably high.
FIG. 1a shows a typical Prior Art (1+3) serpentine cooling design for the first blade of the turbine. The flow path for the 3-pass flow circuit is also shown in FIG. 1b. The airfoil includes a first leading edge cooling passage 2, film cooling holes 8 to deliver cooling air from the leading edge cooling passage 2 to a second leading edge cooling passage 4, a 3-pass serpentine passage having a first leg 20, a second leg 22, and a third leg 24, and trailing edge film cooling passages 40 supplied by cooling air from the first leg 20 of the serpentine passage. Cooling air from the third leg 24 is discharged onto the pressure side and suction side of the blade through pressure side film cooling holes 30 and suction side film cooling holes 32. For a forward flowing 3-pass serpentine cooling design used in the airfoil mid-chord region, the cooling air flows toward and discharges into the high pressure hot gas side pressure section of the pressure side of the blade. In order to satisfy the back flow margin criteria, a high cooling supply pressure is needed in order to prevent the hot gases from flowing into the airfoil.
Since the last leg of the 3-pass serpentine cavities provides film cooling air for both sides of the airfoil, in order to satisfy the back flow margin criteria for the pressure side film row, the internal cavity pressure must be approximately 10% higher than the hot gas pressure of the pressure side of the airfoil. When the cooling air is bled off from the cavity for cooling both the pressure and suction sidewalls, the span-wise internal Mach number becomes lower. This translates to a lower through-flow velocity and lower cooling side internal heat transfer coefficient. The high pressure required preventing inflow from the high pressure side of the airfoil (the pressure side) results in an over-pressuring of the airfoil suction side film holes since the film cooling holes of the pressure side and the suction side is connected to the same cavity.
The U.S. Pat. No. 6,168,381 B1 issued to Reddy on Jan. 2, 2001 and entitled AIRFOIL ISOLATED LEADING EDGE COOLING discloses a serpentine cooling passage design in which an isolation flow chamber (38 in FIG. 3 of this patent) is positioned between a pressure side and suction side flow channels (40 and 42 in FIG. 3), where the pressure side and suction side flow channels are the last leg in a 3-pass serpentine flow circuit, both being supplied with cooling air from a common first and second legs of the 3-pass serpentine circuit. Because both pressure side and suction side flow channels are supplied from the same upstream cooling air passage, the pressures in the pressure and suction side flow channels are the same. The same problem described above exists in the Reddy patent: a high pressure is required to prevent inflow of the hot gasses on the pressure side of the airfoil, and the suction side channel is over-pressurized resulting in excessive flow through the film cooling holes on the suction side of the airfoil.
U.S. Pat. No. 6,595,748 B2 issued to Flodman et al on Jul. 22, 2003 and entitled TRICHANNEL AIRFOIL LEADING EDGE COOLING shows a turbine blade with a 3-pass serpentine flow circuit on the aft end of the blade with a first leg feeding cooling air to the trailing edge discharge holes and a third leg (58 in FIG. 3 of this patent) feeding cooling air to the pressure side film cooling holes and to a suction side cooling channel (second side channel 46 in FIG. 3 of this patent) which feeds cooling air to the suction side of the blade through suction side film cooling holes. A first side channel (44 in FIG. 3 of this patent) feeds cooling air to a shower head arrangement on the leading edge of the blade. The first side channel and the second side channel are on opposite sides of the blade. In the Reddy flow circuit, the suction side channel is still part of the serpentine flow circuit. A metering hole (56 in FIG. 3 of this patent) to feed cooling air to the channel for suction side cooling reduces the pressure and results in lower cooling flow and less cooling. The Flodman patent suffers from the same problem as in the Reddy patent. The supply of cooling air to the suction side of the blade is from a cooling channel (46 in FIG. 3 of this patent) which is delivered through a metering hole (52 in FIG. 3 of this patent). The metering hole reduces the pressure and provides less cooling flow through the cooling channel.